Adiabatic Effectiveness on High-Pressure Turbine Nozzle Guide Vanes Under Realistic Swirling Conditions

Abstract

In modern lean-burn aero-engine combustors, highly swirling flow structures are adopted to control the fuel-air mixing and to provide the correct flame stabilization mechanisms. Aggressive swirl fields and high turbulence intensities are hence expected in the combustor-turbine interface. Moreover, to maximize the engine cycle efficiency, an accurate design of the high-pressure nozzle cooling system must be pursued: in a film-cooled nozzle, the air taken from last compressor stages is ejected through discrete holes drilled on vane surfaces to provide a cold layer between hot gases and turbine components. In this context, the interactions between the swirling combustor outflow and the vane film cooling flows play a major role in the definition of a well-performing cooling scheme, demanding for experimental campaigns at representative flow conditions. An annular three-sector combustor simulator with fully cooled high-pressure vanes has been designed and installed at THT Lab of University of Florence. The test rig is equipped with three axial swirlers, effusion-cooled liners, and six film-cooled high-pressure vanes passages, for a vortex-to-vane count ratio of 1:2. The relative clocking position between swirlers and vanes has been chosen in order to have the leading edge of the central airfoil aligned with the central swirler. In this experimental work, adiabatic film effectiveness measurements have been carried out in the central sector vanes, in order to characterize the film-cooling performance under swirling inflow conditions. The pressure-sensitive paint (PSP) technique, based on heat and mass transfer analogy, has been exploited to catch highly detailed 2D distributions. Carbon dioxide has been used as coolant in order to reach a coolant-to-mainstream density ratio of 1.5. Turbulence and five-hole probe measurements at inlet/outlet of the cascade have been carried out as well, in order to highlight the characteristics of the flow field passing through the cascade and to provide precise boundary conditions. Results have shown a relevant effect of the swirling mainflow on the film cooling behavior. Differences have been found between the central airfoil and the adjacent ones, both in terms of leading edge stagnation point position and of pressure and suction side film coverage characteristics.

Turbulence intensity measurements across a NGV cooled cascade with representative Lean Burn combustor outflow

“…The film cooling scheme is made by 8 rows of cylindrical film cooling holes, with four shower head and four pressure side rows. More details on the combustor simulator rig and on NGVs film cooling system can be found in Bacci et al (2018) and Bacci et al (2018b).…”

“…In addition, a condition without coolant and an in-between condition (M F R = 3.75%) were explored to evaluate the effect of the NGV coolant mass flow rate on the results. A summary of the operating conditions is shown in Table 1, while the row-by-row BR calculation at nominal condition is reported in Bacci et al (2018). It is worth to underline that the exit Mach number of the cascade is relative lower respect to actual engine conditions: even if for instance local sonic conditions could have an impact on the final results, the general conclusions of the paper can be considered valuable to better understand the T u behaviour across cooled NGVs and the impact of cooling flows.…”

“…As stated in the introduction, the results reported in this treatment have been collected by means of five hole probe and hot wire probes traversing on the NGV inlet/outlet planes. Since a thorough characterization of the five hole probe setup can be found in the work of Bacci et al (2018), very little space will be left to its description in this treatment. On the other hand, the setup of HWA measurements, that are the actual goal of this paper, will be deeply detailed in the following.…”

Unsteady Phenomena at the Combustor-Turbine Interface

“…Bacci et al [24,25] used a test rig with cooled vanes and combustion chamber to investigate hot streaks and swirling inflow to the vanes. It was found that different aerofoils face significantly different temperature fields.…”

“…Cubeda et al [27] compared different numerical methods to the experimental measurements of Bacci et al [25,24], finding that neither RANS of combustion chamber and vanes, nor SAS of the combustion chamber used as a boundary condition to steady RANS gave satisfactory results. RANS simulation gave local adiabatic temperature errors exceeding 150K.…”